This thesis details the methods used to build KOTUS (Kit Optimization Tool for Upper Stages) and the application of the tool to the future evolved Ariane 6 upper stage evolution, the Multi-function Upper Stage Express (MUSE), in order to gauge the impact of selected mission-specific kit elements on its payload performance for a range of Earth-bound missions.
Cryogenic bi-propellants such as the liquid hydrogen and oxygen combination are among the most potent propellants, typically sporting specific impulses in excess of 450 s in vacuum. However, the non-propulsive loss of cryogenic propellants sustained in missions that feature multiple boost phases can quickly become prohibitive. These losses limit the efficacy of cryogenic stages for a wide array of mission profiles.
A cryogenic stage is typically optimized for the characteristic challenges imposed by the most prominent mission in the portfolio offered by the launch service provider. The payload performance of this stage is then inevitably suboptimal for the other missions in the portfolio. This performance penalty can be addressed by introducing complementary or supplementary hardware elements into the stage architecture. A new configuration of the stage is then created using a combination of such elements, known as a kit. KOTUS addresses a need to determine the optimal kit for every mission profile at a conceptual level in an expedient manner. This can help to steer the design in the optimal direction from the onset. The tool is capable of performing this analysis for an arbitrarily defined mission, stage, and liquid bi-propellant combination.
In this work, the processes that are relevant to the payload performance analysis are discussed, along with the methods used in the modeling efforts of these processes. This includes the absorbed heat fluxes through insulation, attitude control and propellant orientation, propellant conditioning and pressurization, and feed line management. The agglomeration of these processes allow the modeling of the performance of a baseline stage when embarking on any mission. A set of elements was subsequently curated that can be equipped either individually or in combination with others to this baseline stage. The elements chosen are primarily focused on enhancing the operational performance of the stage by reducing non-propulsive losses. The integrated element set includes various insulation options, boost pumps, feed line heat exchangers, vapor-cooled shields, and thermodynamic venting systems. The modeling efforts for each of these elements and their individual impact on the upper stage relative to the baseline configuration is discussed at length. Sensitivity studies of uncertain input parameters and correction factors were conducted, as well as sub-model validations, to provide an indication of the tool’s credibility.
Results showed that for medium duration missions with one or more restarts, payload performance enhancements upwards of 100 kg are possible when a deployable sun shield is added to the baseline configuration. The degree to which the sun shield is able to limit absorbed heat flows allowed it to drastically reduce non-propulsive losses. For single boost missions, only the limiting of thermal residuals was demonstrated to be relevant to enhancing the payload performance of the stage. This was shown to be best accomplished through the use of a feed line heat exchanger in the LH2 feed line and a turbine-driven boost pump in the LOX feed line, the combination of which is estimated to enable an extra payload mass of 103 kg. In addition, it was shown that the default pressurization setup of the reference stage (i.e., the heterogeneous pressurization of the LOX tank and autogenous pressurization of the LH2 tank) was already optimal, and that no tank or feed line sizing changes would be beneficial.
The potential for the deployment of hardware kits has been demonstrated. Even short duration, single boost missions were shown to be able to benefit from tailor-made kit configurations, though the potential for payload performance enhancement is significantly larger for medium to long duration missions, particularly those featuring multiple engine restarts. A number of future work recommendations are listed in the conclusion, chief among them to verify the assumptions made in this work and to improve the determination of thermal residuals through enhanced knowledge of incident heat fluxes and spatial energy distributions in the propellant.